Barrier to prevent super alloy depletion into nickel-cbn blade tip coating

ABSTRACT

A diffusion barrier coating on a nickel-based alloy substrate comprising the diffusion barrier being coupled to the substrate between the substrate and a abrasive composite material opposite the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material a high phosphorus nickel-P alloy.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation-in-part of U.S. application Ser. No. 16/720,847 filed Dec. 19, 2019.

BACKGROUND

The present disclosure is directed to a diffusion barrier layer for integrally bladed rotor tip Nickel-Cubic Boron Nitride (Ni-CBN) plating.

In certain gas turbine engines, the nickel integrally bladed rotor is suffering lost life time of the tip Ni-CBN coating. Elements of the base super alloy diffuse from the base super alloy into the Ni-CBN layer after engine run or heat treatment. Elements such as Cr and Al diffuse from the base super alloy into the Ni-CBN coating layer.

As a result of the diffusion of the elements from the base super alloy and the propensity of these elements to oxidize during engine operation, oxides form along surfaces and grain boundaries within the coating. These oxides reduce the strength of the coating causing loss of CBN particles and recession of the coating.

What is needed is a technique to diminish the diffusion and subsequent nickel alloy depletion.

SUMMARY

In accordance with the present disclosure, there is provided a diffusion barrier coating on a nickel-based alloy substrate comprising the diffusion barrier coupled to the substrate between the substrate and a composite material opposite the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material or a high phosphorus nickel-P alloy.

In another and alternative embodiment, the nickel phosphorus alloy material comprise a layered coating structure.

In another and alternative embodiment, the diffusion barrier consists of plated layers.

In another and alternative embodiment, the layered coating includes multiple layers.

In another and alternative embodiment, the diffusion barrier comprises a nickel strike layer between the substrate and the diffusion barrier.

In accordance with the present disclosure, there is provided a gas turbine engine component comprising a compressor integrally bladed rotor having a blade with an airfoil section and a tip having a substrate; a diffusion barrier coupled to the substrate between the substrate and a composite material opposite the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material or a high phosphorus nickel-P alloy.

In another and alternative embodiment, the diffusion barrier includes multiple layers.

In another and alternative embodiment, the diffusion barrier comprises a nickel strike layer between the substrate and the diffusion barrier.

In another and alternative embodiment, the substrate comprises a nickel-based alloy.

In another and alternative embodiment, the integrally bladed rotor is located in a high pressure compressor section of the gas turbine engine.

In accordance with the present disclosure, there is provided a process for diffusion inhibition in a nickel-based alloy substrate of a gas turbine engine component comprising applying a diffusion barrier coupled to the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material or a high phosphorus nickel-P alloy; and subjecting the gas turbine engine component with nickel-based alloy substrate to at least one of a heat treatment and an engine operation.

In another and alternative embodiment, the diffusion barrier includes multiple layers.

In another and alternative embodiment, the process further comprises preventing Cr, Al, and Ti depletion from the nickel-based alloy substrate by reducing diffusion between the nickel-based alloy substrate and the diffusion barrier.

Other details of the diffusion barrier are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a simplified cross-sectional view of a gas turbine engine.

FIG. 2 is a cross sectional schematic of an exemplary coating system.

DETAILED DESCRIPTION

FIG. 1 is a simplified cross-sectional view of a gas turbine engine 10 in accordance with embodiments of the present disclosure. Turbine engine 10 includes fan 12 positioned in bypass duct 14. Turbine engine 10 also includes compressor section 16, combustor (or combustors) 18, and turbine section 20 arranged in a flow series with upstream inlet 22 and downstream exhaust 24. During the operation of turbine engine 10, incoming airflow Fi enters inlet 22 and divides into core flow Fc and bypass flow FB, downstream of fan 12. Core flow Fc continues along the core flowpath through compressor section 16, combustor 18, and turbine section 20, and bypass flow FB proceeds along the bypass flowpath through bypass duct 14.

Compressor 16 includes stages of compressor vanes 26 and blades 28 arranged in low pressure compressor (LPC) section 30 and high pressure compressor (HPC) section 32. Turbine section 20 includes stages of turbine vanes 34 and turbine blades 36 arranged in high pressure turbine (HPT) section 38 and low pressure turbine (LPT) section 40. HPT section 38 is coupled to HPC section 32 via HPT shaft 42, forming the high pressure spool. LPT section 40 is coupled to LPC section 30 and fan 12 via LPT shaft 44, forming the low pressure spool. HPT shaft 42 and LPT shaft 44 are typically coaxially mounted, with the high and low pressure spools independently rotating about turbine axis (centerline) CL.

Combustion gas exits combustor 18 and enters HPT section 38 of turbine 20, encountering turbine vanes 34 and turbines blades 36. Turbine vanes 34 turn and accelerate the flow of combustion gas, and turbine blades 36 generate lift for conversion to rotational energy via HPT shaft 42, driving HPC section 32 of compressor 16. Partially expanded combustion gas flows from HPT section 38 to LPT section 40, driving LPC section 30 and fan 12 via LPT shaft 44. Exhaust flow exits LPT section 40 and turbine engine 10 via exhaust nozzle 24. In this manner, the thermodynamic efficiency of turbine engine 10 is tied to the overall pressure ratio (OPR), as defined between the delivery pressure at inlet 22 and the compressed air pressure entering combustor 18 from compressor section 16. As discussed above, a higher OPR offers increased efficiency and improved performance. It will be appreciated that various other types of turbine engines can be used in accordance with the embodiments of the present disclosure.

Referring now to FIG. 2 , there is illustrated a turbine engine component 50, such as a compressor integrally bladed rotor or blade or vane. The component 50 can be an integrally bladed rotor in the high pressure compressor section 32 of the gas turbine engine 10 The turbine engine component 50 has an airfoil portion 52 with a tip 54.

The turbine engine component 50 may be formed from a titanium-based alloy or a nickel-based alloy. On the substrate tip 54 of the airfoil portion 52, a abrasive material 56 is applied for rub and abradability against an abradable coating (not shown). In an exemplary embodiment the abrasive material 56 can be a nickel-cubic boron nitride (Ni-CBN) material.

A diffusion barrier 58 can be coupled to the tip substrate 54 between the tip substrate 52 and the abrasive material 56. In an exemplary embodiment, the diffusion barrier 58 comprises an electrolytic nickel cobalt with a chromium aluminum yttria powder (Ni—Co with Cr—Al—Y) coating or a high phosphorus nickel-P alloy. In an exemplary embodiment, the diffusion barrier 58 including Cr—Al—Y powders embedded Ni—Co electro-plating layer or a high phosphorus nickel-P alloy. 58 can act as a bond coat.

In an exemplary embodiment, the diffusion barrier coating 58 can replace the traditional columnar structure of prior bond coat. In an exemplary embodiment, the diffusion barrier coating 58 can replace the traditional unalloyed Ni of prior bond coat. The addition of alloying elements (esp. Al, Y, Cr or P) to the diffusion barrier 58 reduces the chemical potential for diffusion of these elements from the blade tip 52.

The inclusion of the Cr—Al—Y embedded Ni—Co or a high phosphorus nickel-P alloy eliminates the Cr diffusion from the nickel alloy substrate 54. In an exemplary embodiment, aluminum depletion occurs from the Cr—Al—Y thus forming Ni alloy layer under the grits 62.

In an exemplary embodiment, the diffusion barrier 58 can include a high phosphorus nickel-P alloy layer 66 on top of the nickel strike layer 70. A nickel strike layer 70 can be applied to the tip substrate 52.

A technical advantage of the diffusion barrier is that it prevents Cr, Al, and Ti depletion from the base alloy of the substrate.

Another technical advantage of the diffusion barrier includes very low grain boundary oxidation.

Another technical advantage of the disclosed diffusion barrier includes prevention of the Ni super alloy depletion after engine operation.

Another technical advantage of the disclosed diffusion barrier includes elimination of potential mechanical strength reduction due to the depletion of the alloy chemistry.

Another technical advantage of the disclosed diffusion barrier includes extending the lifetime of the IBR used in the HPC section.

There has been provided a diffusion barrier. While the diffusion barrier has been described in the context of specific embodiments thereof, other unforeseen alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations which fall within the broad scope of the appended claims. 

What is claimed is:
 1. A diffusion barrier coating on a nickel-based alloy substrate comprising: the diffusion barrier coupled to the substrate between the substrate and a abrasive material opposite the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material or a high phosphorus nickel-P alloy.
 2. The diffusion barrier coating on a substrate according to claim 1, wherein said nickel-P alloy comprises a high phosphorus layered grain structure.
 3. The diffusion barrier coating on a substrate according to claim 2, wherein said diffusion barrier consists of a nickel cobalt and chromium powder material.
 4. The diffusion barrier coating on a substrate according to claim 1, wherein said abrasive material comprises a nickel-cubic boron nitride material.
 5. The diffusion barrier coating on a substrate according to claim 1, wherein said diffusion barrier comprises a bond coat between said substrate and said abrasive material.
 6. The diffusion barrier coating on a substrate according to claim 1, wherein said diffusion barrier comprises a nickel strike layer between said substrate and said diffusion barrier.
 7. A gas turbine engine component comprising: a compressor integrally bladed rotor having a blade with an airfoil section and a tip having a substrate; a diffusion barrier coupled to the substrate between the substrate and a composite material opposite the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material or a high phosphorus nickel-P alloy.
 8. The gas turbine engine component according to claim 7, wherein said nickel cobalt and chromium-aluminum-yttria powder material comprises a bond layer.
 9. The gas turbine engine component according to claim 7, wherein said high phosphorus nickel-P alloy comprises a layered grain structure.
 10. The gas turbine engine component according to claim 9, wherein the diffusion barrier includes multiple layers.
 11. The gas turbine engine component according to claim 9, wherein said diffusion barrier comprises a nickel strike layer between said substrate and said diffusion barrier.
 12. The gas turbine engine component according to claim 7, wherein said substrate comprises a nickel-based alloy.
 13. The gas turbine engine component according to claim 7, wherein said integrally bladed rotor is located in a high pressure compressor section of the gas turbine engine.
 14. A process for diffusion inhibition in a nickel-based alloy substrate of a gas turbine engine component comprising: applying a diffusion barrier coupled to the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material or a high phosphorus nickel-P alloy; coating said diffusion barrier under a abrasive composite; and subjecting said gas turbine engine component with nickel-based alloy substrate to at least one of a heat treatment and an engine operation.
 15. The process of claim 14, coating said a nickel cobalt and chromium-aluminum-yttria powder material coating or a high phosphorus nickel-P alloy materials comprises a bond coat.
 16. The process of claim 14, wherein the diffusion barrier includes multiple layers.
 17. The process of claim 14, wherein said matrix abrasive material comprises a nickel-cubic boron nitride material.
 18. The process of claim 14, further comprising: preventing Cr, Al, and Ti depletion from the nickel-based alloy substrate by reducing diffusion between said nickel-based alloy substrate and said matrix abrasive composite with said diffusion barrier. 